Gas turbine engines are well known in the art. FIG. 1 shows a known ducted fan gas turbine engine 10 having a principal axis of rotation 11 and comprising, in axial flow series: an air intake 12, a propulsive fan 14, an intermediate pressure compressor 16, a high-pressure compressor 18, a combustor 20, a high-pressure turbine 22, an intermediate pressure turbine 24, a low-pressure turbine 26 and a core exhaust nozzle 28. A nacelle 30 generally surrounds the engine 10 and defines the intake 12, a bypass duct 32 and a bypass exhaust nozzle 34. It may also include a thrust reverser 36.
Air entering the intake 12 is accelerated by the fan 14 to produce a bypass flow and a core flow. The bypass flow travels down the bypass duct and exits the bypass exhaust nozzle to provide the majority of the propulsive thrust produced by the engine 10. The core flow enters in series the intermediate pressure compressor 16, high pressure compressor 18 and the combustor 20, where fuel is added to the compressed air and the mixture burned. The hot combustion products expand through and drive the high, intermediate and low-pressure turbines 22, 24, 26 before being exhausted through the nozzle 28 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines 22, 24, 26 respectively drive the high and intermediate pressure compressors 18, 16 and the fan 14 by suitable interconnecting shafts.
If the bypass or cold nozzle 34 is upstream of the core exhaust or hot nozzle 28 then the engine may be referred to as having separate jets. If the bypass exhaust nozzle 34 extends aft of the core exhaust nozzle 28 and encloses it, then the engine 10 is said to have a mixed exhaust. In that case the bypass or final nozzle is often referred to as a mixed or common nozzle.
Variable area mixed flow exhaust nozzles are widely used on military turbofan engines. Most variable area cold nozzle designs work by varying the outside diameter of the nozzle, but some designs change the inside diameter of the nozzle by means of a variable geometry afterbody as described for example in the U.S. Pat. No. 3,756,026.
The benefit of having a variable area cold nozzle is for controlling the working line of the fan for improved fan efficiency, surge margin and operability particularly in turbofan engines with low pressure ratio fans or reheat systems.
The working line is the locus of fan pressure ratio plotted against fan inlet non-dimensional flow (or mass flow corrected to standard pressure and temperature) for normal steady-state engine operation. At any non-dimensional inlet flow there is an optimum fan pressure ratio for highest efficiency and an upper limit for fan pressure ratio, beyond which the streamline flow through the fan will break down and the fan will surge. At all conditions the non-dimensional flows are determined by the effective fan nozzle exit area and nozzle pressure ratio.
When the fan pressure ratio is high or the engine is flying at high subsonic (or supersonic) Mach number, the final nozzle will have sonic or near sonic flow and is said to be choked. Under these conditions the fan will have a steep working line which will tend to track the locus of peak fan efficiency and run parallel to the surge line as power is varied, providing a safe margin with respect to surge. However, if the fan pressure ratio and the airspeed are lower, the flow through the nozzle will be subsonic and the non-dimensional flow through the nozzle will reduce with reducing fan power and fan and nozzle pressure ratios.
In this case the non-dimensional mass flow at entry to the fan will reduce more rapidly as fan speed and power are reduced and so the fan working line will be flatter. This means that the working line no longer follows the locus of peak efficiency and could be too high at low power conditions where the fan may now surge. Conversely, if a larger fan exit nozzle area is provided, the fan efficiency will suffer when the engine is operated at high airspeeds, such as at cruise at altitude, because here the working line will be too low. These problems become more severe as a fan is designed for lower pressure ratios, below about 1.45, and in this case a variable area nozzle can significantly improve fan efficiency at cruise and top of climb conditions.
An alternative design using a mixed-flow final nozzle of fixed geometry to improve the fan working line is used on several Rolls-Royce engines such as the Trent 700. In this arrangement the core exhaust and the fan bypass section exhaust are admitted into a common duct and share a common final exhaust nozzle. As the engine is throttled back the core exhaust mass flow reduces more rapidly than the fan bypass section mass flow and occupies a smaller proportion of the final nozzle cross-section, increasing the effective flow area available to the fan bypass flow. This arrangement is helpful, but ultimately not as effective as a variable area nozzle, because it mostly only responds to changes in fan pressure ratio or power level and not to changes in flight speed.
U.S. Pat. No. 6,070,407 describes a bypass duct of a gas turbine engine which is provided with a secondary duct at least partly within the downstream end of the bypass duct. The secondary duct is provided with means such as flaps whereby the airflow therethrough may be varied to suit the flight requirements of an associated aircraft, in a way which will control the maximum diameter of the free stream tube airflow at the intake of the engine, thus effectively reducing the frontal area of the fan duct, and therefore, drag.
A similar arrangement to that described in U.S. Pat. No. 6,070,407 is described in US2008302083, but for an entirely separate purpose. Here, the described aircraft has at least one turbofan engine assembly having a shrouded core engine, a short outer nacelle surrounding a fan and a forward portion of the core engine, and a fan exhaust duct through the nacelle. A mixer duct shell is positioned coaxially with the engine shroud and extends forwardly into the fan duct to provide an interstitial mixer duct between the mixer duct shell and the core engine shroud. The aft portion of the mixer duct shell extends over a turbine exhaust frame, an attached mixer (if included), and a tail cone exhaust plug. The mixer duct shell is described as reducing noise and plume exhaust heat radiated from aircraft turbofan engines.
A forward portion of the shell in US′083 is affixed to the core engine shroud by a plurality of circumferentially spaced and aerodynamically tailored radial pillars. An aft portion of the shell may be moved in an aft-ward direction along a pillar slide with weight supported on a sliding track attached to the engine pylon sidewall. Moving the aft portion provides access to underlying structure and is carried out only whilst the engine is not operating. When operating, the aft portion is locked to the core engine in a fixed relation.
GB1207194 describes a jet engine arranged for the suppression of jet sound and comprises nacelle surrounding a duct for exhaust gas flow; flaps positioned to form a converging section adjacent the end of the duct; blow-in doors pivotally mounted at the end of the nacelle arranged to move inwardly; an annular body positioned rearwardly of the pod or nacelle with the inner surface of the body forming an inlet passageway with each blow-in door, and turbulence inducing means to produce a shear layer surrounding the expanding exhaust flow.
The present invention seeks to provide an improved variable area nozzle arrangement for a gas turbine engine.